A combustion section of a gas turbine generally includes a plurality of combustors that are arranged in an annular array around an outer casing such as a compressor discharge casing. Pressurized air flows from a compressor to the compressor discharge casing and is routed to each combustor. Fuel from a fuel nozzle is mixed with the pressurized air in each combustor to form a combustible mixture within a primary combustion zone of the combustor. The combustible mixture is burned to produce hot combustion gases having a high pressure and high velocity. The combustion gases are routed towards an inlet of a turbine of the gas turbine through a hot gas path that is at least partially defined by an annular combustion liner and/or an annular transition duct. The hot gas path extends through the turbine and terminates at an outlet of the turbine.
The constant demand for increased operating temperatures in gas turbine engines has necessitated the development of various coating materials such as ceramics that can be applied to the various hot gas paths components such as the combustion liner, the transition duct and/or turbine nozzles and turbine blades to insulate those components from the heat contained in the combustion gases, thereby extending the life of those components. These coatings are known in the art as thermal barrier coatings (TBC).
A thermal barrier coating typically comprises at least one layer of a refractory or thermally insulating material having a low thermal conductivity such as about 1-3 W/(m)(K). The coating material may be applied by one of known deposition techniques such as a thermal or plasma spray process or a physical vapor deposition process. Typically, a thermal barrier coating is applied in multiple layers. In particular applications, a bond-coat is applied to an inner or hot side surface of the liner or transition duct. The bond-coat provides a layer which adheres well to the underlying alloy and that provides protection against oxidation of the alloy. The refractory or thermal insulation coat is then applied over the bond-coat. Some thermal barrier coatings may also include an intermediate layer or interlayer that is applied over the bond-coat. The interlayer may provide improved adhesion for the final thermal insulating coat.
Despite great care taken during manufacture to ensure adhesion of the thermal insulation coat to the bond-coat, thermal cycling eventually leads to subsurface defects in the thermal barrier coating known as delamination or disbonding. Delamination generally leads to spallation that eventually exposes the underlying alloy to extreme temperatures that may impact the durability of the liner and/or transition duct. As a result, the thermal barrier coating must be inspected for subsurface defects during scheduled maintenance or planned outages of the gas turbine. Typically, the thermal barrier coating is stripped and replaced after a pre-determined number of inspection cycles, in part due to limitations of many current inspection processes.
Removal and reapplication of the thermal barrier coating significantly increases the time required to inspect the hot gas path component and contributes substantially to the overall cost of inspection/repair. Therefore, there is a continued need to provide a non-destructive method for determining size and location of subsurface defects in the thermal barrier coating to qualify the risk associated with continued use of the thermal barrier coating.